1. Field of the Invention
The present invention relates to multi-stage launch vehicles. In particular, the present invention relates to relatively high performance launch vehicles for placing payloads into earth orbit and beyond, i.e., escape from the gravity of Earth.
2. Description of Prior Art and Related Information
The various approaches to launch vehicle design may be generally classified into single stage or multistage launch vehicle systems. Single stage launch vehicles employ a single thruster stage which includes all the propellant required to deliver a specified velocity to the payload. Since considerable mass is contained in the propellant tanks, engines and thrust structure, which mass becomes unnecessary once propellant therein is expended, a single stage launch vehicle is inherently of less than optimum efficiency. Multi-stage launch vehicles, where an entire stage, including propellant tanks and engines, is jettisoned after expenditure, have accordingly been developed and gained predominance for earth orbit launch applications. Due to their simplicity, however, single stage launch vehicles will, in general, be cheaper and more reliable than multistage launch vehicles.
A single stage vehicle with multiple engines which can be staged has also been employed. This vehicle is referred to as the Atlas launch vehicle and is made by General Dynamics. The Atlas stage is operated in a mode where only engines and part of their thrust structure are staged at an appropriate time in flight. (Due to this partial staging, this is commonly called a stage and a half launch vehicle.) A stage and a half launch vehicle has two important features: (1) it reduces total weight at a time in flight when the jettisoned weight is no longer necessary to the efficient performance of the stage, and (2) it reduces the thrust at a time when the propellant weight has been reduced to the point that, with all engines continuing to thrust, the acceleration loads (thrust to weight ratio) delivered to the stage and its payload would be greater than desired from a structural design standpoint.
The Atlas was originally designed as an ICBM (Intercontinental Ballistic Missile). The velocity requirements for ICBMs are substantially lower than the velocities required to place a payload into Earth Orbit. To achieve increased capabilities for this launch vehicle updating of engine thrusts and lengthening of the stage(s) to accommodate increases in propellant have been employed. Substantial further increases in performances of these vehicles are inhibited, however, by the difficulties associated with achieving further increases in engine thrusts and increased stage lengths. In particular, the stage length problem is very severe since the ratio of the overall length to diameter of a launch vehicle is critical to its stiffness, which in turn, is critical to the dynamic loads it can withstand due to high altitude winds that it encounters as it traverses the Earth's atmosphere. Also, the wind loads increase as the length of the vehicle increases. Very similar problems are also presented with providing increased payload to Earth Orbit capabilities for the other expandable U.S. launch vehicles, the Titan and Delta launch vehicles, since these were also originally designed as ICBMs or IRBMs (Intermediate Range Ballistic Missiles).
These difficulties in providing further upgrades in capability may be appreciated by consideration of one specific upgraded Atlas launch vehicle, the Atlas IIA. This upgraded Atlas, the Atlas IIA is one of three current unmanned expendable U.S. Space Launch Vehicles. The others are the Titan IV and Delta II. An Atlas IIA configuration is shown in FIG. 1. The Atlas IIA is a two and a half stage launch vehicle with the lower stage being an Atlas one and a half stage having three engines 1, 2, and 3, two of which, 1 and 3, are booster engines which are staged during first burn and have a thrust of about 270,000 lbs. The sustainer engine 2 which continues to thrust until the Atlas lower stage burns out, has a smaller thrust (about 81,000 lbs.) than the booster engines 1 and 3. The upper stage 4 is a Centaur rocket, also built by General Dynamics. The Centaur is propelled by two LO.sup.2 /H.sup.2 cryogenic engines. The particular Atlas IIA configuration shown in FIG. 1 has a payload performance capability to Geosynchronous Transfer Orbit (GTO) of about 6200 lbs and an overall length L of 155 ft. The Atlas II family of launch vehicles based on the general Atlas/Centaur series stages, has several configurations which range in payload capability to GTO from about 3800 lbs. to 8200 lbs.
The Atlas lower stage, has undergone several design changes to improve its performance for space launch missions. However, increased engine thrusts have reached a level where quantum improvements will require new engine developments. Also, increased velocity capabilities of the stage have been achieved by lengthening the stage to accommodate more propellant while its original ten foot diameter has remained fixed. As mentioned above, the length to diameter ratio of a launch vehicle is critical to its stiffness, which in turn is critical to the dynamic loads it can withstand due to high altitude winds it encounters as it traverses the Earth's atmosphere. The wind loads must be limited both to protect the structural integrity of the vehicle and to maintain its control authority by means of the engines thrusting at varying gimbal angles to maintain the proper vehicle attitude. Small increases in performance have been achieved by employing load alleviating devices which reduce the wind loads. Although stiffening the Atlas is a possibility, the design changes required to achieve a substantial improvement would be extensive. However, it is clear that large performance increases will not be achieved without substantial increases in the stiffness of the vehicles and increases in engine thrust either by new engine developments or increasing the numbers of existing engines in the vehicle.
The ultimate effect of this problem is to limit Atlas II launches to specified high altitude wind conditions. This, in turn, limits the launch availability of the vehicle at a time when high launch vehicle availability is becoming a primary operational requirement.
Launch vehicles utilizing Atlas stages burning in parallel might be employed to overcome the length to diameter ratio barrier to Atlas II growth. FIGS. 2(b) and 2(c) illustrate two launch vehicles utilizing one and two Atlas stage "strap ons", respectively, to a modified Atlas IIA stage with a Centaur upper stage. The Atlas IIA configuration illustrated in FIG. 1 is also shown for comparison in FIG. 2(a). A modified Atlas II launch vehicle 5 with a single Atlas stage "strap on" would have an overall length L somewhat greater than the Atlas IIA, i.e., L"=approximately 180 ft. This is due to the fact that the Atlas "strap on" stage 6 and the Atlas core stage 7, optimized for performance, are about 0.75 and 1.2 times the length of the conventional Atlas IIA stage shown in FIG. 2(a). Thus, the vehicle 5 does not present an attractive option for growth because the overall vehicle length to diameter ratio would require substantial redesign to deal with the high altitude wind loads problem. Another modified launch vehicle 8, employing an Atlas IIA with a long Atlas stage 9 and Centaur upper stage with two additional Atlas stage strap ons, shown in FIG. 2(c), would have an even greater overall length, L"=186 ft. An additional consideration for these two launch vehicle configurations is that the first stage would be staged before reaching the high altitude winds regime that would cause load problems for the long Atlas second stage 7, 9 plus Centaur stage plus payload.
In addition to the above noted problems, current U.S. launch vehicles are also limited by their reliability. All of the engines in the Atlas IIA, Titan IV and Delta II must function properly to achieve a successful launch. Accordingly, most launch vehicle failures are due to one of the engine's failing. This problem cannot be dealt with by adopting an engine-out strategy because none of these vehicles have a sufficient number of engines to meet their mission performance requirements with an engine failed (engine-out capability).
Accordingly, there presently exists a need to improve the performance, reliability and cost effectiveness of one or more of the current U.S. expendable launch vehicles.